Repair procedure for a gas turbine engine via variable polarity welding

ABSTRACT

A liner panel for a gas turbine engine includes a cold side having a backside feature and a hot side with a variable polarity weld.

BACKGROUND

The present disclosure relates to a gas turbine engine and, more particularly, to a combustor section therefor.

Gas turbine engines, such as those that power modern commercial and military aircraft, generally include a compressor to pressurize an airflow, a combustor for burning a hydrocarbon fuel in the presence of the pressurized air, and a turbine to extract energy from the resultant combustion gases.

Combustors are subject to high thermal loads for prolonged periods of time. To alleviate the accompanying thermal stresses, the combustor liners are cooled and include various thermal barrier coatings. In one cooling arrangement a twin wall configuration includes a shell lined with heat shields often referred to as Impingement Film-Cooled Floatwall (IFF) liner panels which are attached to the shell with studs and nuts. Dilution holes in the shell are aligned with respective dilution holes in the liner for introduction of dilution air. In addition to the dilution holes, relatively smaller air impingement holes direct cooling air between the outer shell and the liner panels to cool the backside of the liner panels. This cooling air then exits effusion holes in the liner panels to form a cooling film on the hot side of the liner panels.

Although operationally effective, combustor float wall panels may, over time, exhibit cracking, erosion, and oxidation damage that is not economically repairable as the panels are relatively thin and contain features that may be damaged or distorted. Additionally, the protective coating cannot be removed from the features without unacceptable levels of material loss.

SUMMARY

A liner panel for use in a combustor of a gas turbine engine according to one disclosed non-limiting embodiment of the present disclosure includes a cold side having a backside feature and a hot side with a variable polarity weld.

In a further embodiment of the foregoing embodiment, the hot side includes a thermal barrier coating. In the alternative or additionally thereto, in the foregoing embodiment the thermal barrier coating is a nickel-aluminide. In the alternative or additionally thereto, in the foregoing embodiment the thermal barrier coating is an aluminum oxide.

In a further embodiment of any of the foregoing embodiments, the cold side includes a coating. In the alternative or additionally thereto, in the foregoing embodiment the coating is a diffusion-aluminide.

A method of repairing a component of a gas turbine engine according to another disclosed non-limiting embodiment of the present disclosure includes welding a damaged area via variable polarity welding.

In a further embodiment of the foregoing embodiment, the method further comprising stripping a thermal barrier coating prior to the welding. In the alternative or additionally thereto, the foregoing embodiment further comprising applying a thermal barrier coating after the welding.

In a further embodiment of any of the foregoing embodiments, the method further comprising acid stripping a thermal barrier coating prior to the welding.

In a further embodiment of any of the foregoing embodiments, the method further comprising stripping a thermal barrier coating in the damaged area prior to the welding. In the alternative or additionally thereto, the foregoing embodiment, further comprising applying a thermal barrier coating in the damaged area after the welding.

A liner panel for use in a combustor of a gas turbine engine according to another disclosed non-limiting embodiment of the present disclosure includes a cold side having a coating and a hot side with a variable polarity weld.

In a further embodiment of the foregoing embodiment, the hot side includes a thermal barrier coating. In the alternative or additionally thereto, in the foregoing embodiment the thermal barrier coating is a nickel-aluminide.

In a further embodiment of any of the foregoing embodiments, the cold side includes a coating. In the alternative or additionally thereto, in the foregoing embodiment the coating is a diffusion-aluminide.

BRIEF DESCRIPTION OF THE DRAWINGS

Various features will become apparent to those skilled in the art from the following detailed description of the disclosed non-limiting embodiment. The drawings that accompany the detailed description can be briefly described as follows:

FIG. 1 is a schematic cross-section of a gas turbine engine;

FIG. 2 is a partial longitudinal schematic sectional view of a combustor section according to one non-limiting embodiment that may be used with the gas turbine engine shown in FIG. 1;

FIG. 3 is an exploded view of a liner panel of the combustor; and

FIG. 4 is a schematic block diagram of a repair process.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gas turbine engine 20 is disclosed herein as a two-spool turbo fan that generally incorporates a fan section 22, a compressor section 24, a combustor section 26 and a turbine section 28. Alternative engines might include an augmentor section (not shown) among other systems or features. The fan section 22 drives air along a bypass flowpath while the compressor section 24 drives air along a core flowpath for compression and communication into the combustor section 26 then expansion through the turbine section 28. Although depicted as a turbofan in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with turbofans as the teachings may be applied to other types of turbine engines such as a turbojets, turboshafts, and three-spool (plus fan) turbofans wherein an intermediate spool includes an intermediate pressure compressor (“IPC”) between a Low Pressure Compressor (“LPC”) and a High Pressure Compressor (“HPC”), and an intermediate pressure turbine (“IPT”) between the high pressure turbine (“HPT”) and the Low pressure Turbine (“LPT”).

The engine 20 generally includes a low spool 30 and a high spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine static structure 36 via several bearing structures 38. The low spool 30 generally includes an inner shaft 40 that interconnects a fan 42, a low pressure compressor 44 (“LPC”) and a low pressure turbine 46 (“LPT”). The inner shaft 40 drives the fan 42 directly or through a geared architecture 48 to drive the fan 42 at a lower speed than the low spool 30. An exemplary reduction transmission is an epicyclic transmission, namely a planetary or star gear system.

The high spool 32 includes an outer shaft 50 that interconnects a high pressure compressor 52 (“HPC”) and high pressure turbine 54 (“HPT”). A combustor 56 is arranged between the high pressure compressor 52 and the high pressure turbine 54. The inner shaft 40 and the outer shaft 50 are concentric and rotate about the engine central longitudinal axis A which is collinear with their longitudinal axes.

Core airflow is compressed by the LPC 44 then the HPC 52, mixed with the fuel and burned in the combustor 56, then expanded over the HPT 54 and the LPT 46. The turbines 54, 46 rotationally drive the respective low spool 30 and high spool 32 in response to the expansion. The main engine shafts 40, 50 are supported at a plurality of points by bearing structures 38 within the static structure 36. It should be understood that various bearing structures 38 at various locations may alternatively or additionally be provided.

In one non-limiting example, the gas turbine engine 20 is a high-bypass geared aircraft engine. In a further example, the gas turbine engine 20 bypass ratio is greater than about six (6:1). The geared architecture 48 can include an epicyclic gear train, such as a planetary gear system or other gear system. The example epicyclic gear train has a gear reduction ratio of greater than about 2.3, and in another example is greater than about 2.5:1. The geared turbofan enables operation of the low spool 30 at higher speeds which can increase the operational efficiency of the low pressure compressor 44 and low pressure turbine 46 and render increased pressure in a fewer number of stages.

A pressure ratio associated with the low pressure turbine 46 is pressure measured prior to the inlet of the low pressure turbine 46 as related to the pressure at the outlet of the low pressure turbine 46 prior to an exhaust nozzle of the gas turbine engine 20. In one non-limiting embodiment, the bypass ratio of the gas turbine engine 20 is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor 44, and the low pressure turbine 46 has a pressure ratio that is greater than about five (5:1). It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present disclosure is applicable to other gas turbine engines including direct drive turbofans.

In one embodiment, a significant amount of thrust is provided by the bypass flow path B due to the high bypass ratio. The fan section 22 of the gas turbine engine 20 is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet. This flight condition, with the gas turbine engine 20 at its best fuel consumption, is also known as bucket cruise Thrust Specific Fuel Consumption (TSFC). TSFC is an industry standard parameter of fuel consumption per unit of thrust.

Fan Pressure Ratio is the pressure ratio across a blade of the fan section 22 without the use of a Fan Exit Guide Vane system. The low Fan Pressure Ratio according to one non-limiting embodiment of the example gas turbine engine 20 is less than 1.45. Low Corrected Fan Tip Speed is the actual fan tip speed divided by an industry standard temperature correction of (“T”/518.7^(0.5)) in which “T” represents the ambient temperature in degrees Rankine. The Low Corrected Fan Tip Speed according to one non-limiting embodiment of the example gas turbine engine 20 is less than about 1150 fps (351 m/s).

With reference to FIG. 2, the combustor 56 generally includes an outer combustor liner assembly 60, an inner combustor liner assembly 62 and a diffuser case module 64. The outer combustor liner assembly 60 and the inner combustor liner assembly 62 are spaced apart such that a combustion chamber 66 is defined therebetween. The combustion chamber 66 is generally annular in shape. The outer combustor liner assembly 60 is spaced radially inward from an outer diffuser case 64-O of the diffuser case module 64 to define an outer annular plenum 76. The inner combustor liner assembly 62 is spaced radially outward from an inner diffuser case 64-I of the diffuser case module 64 to define an inner annular plenum 78. It should be understood that although a particular combustor is illustrated, other combustor types with various combustor liner arrangements will also benefit herefrom. It should be further understood that the disclosed cooling flow paths are but an illustrated embodiment and should not be limited only thereto.

The combustor liner assemblies 60, 62 contain the combustion products for direction toward the turbine section 28. Each combustor liner assembly 60, 62 generally includes a respective support shell 68, 70 which supports one or more liner panels 72, 74 mounted to a hot side of the respective support shell 68, 70. The liner panels 72, 74, often referred to as Impingement Film Float (IFF) wall panels or heat shields, define an array which form the annular combustor chamber 66. Each of the liner panels 72, 74 may be generally rectilinear and manufactured of, for example, a nickel based super alloy, ceramic or other temperature resistant material. In one disclosed non-limiting embodiment, the array includes a multiple of forward liner panels 72A and a multiple of aft liner panels 72B that line the hot side of the outer shell 68 and a multiple of forward liner panels 74A and a multiple of aft liner panels 74B that line the hot side of the inner shell 70.

The combustor 56 further includes a forward assembly 80 immediately downstream of the compressor section 24 to receive compressed airflow therefrom. The forward assembly 80 generally includes an annular hood 82, a bulkhead assembly 84, a multiple of fuel nozzles 86 (one shown) and a multiple of fuel nozzle guides 90 (one shown). Each of the fuel nozzle guides 90 is circumferentially aligned with one of the hood ports 94 to project through the bulkhead assembly 84. Each bulkhead assembly 84 includes a bulkhead support shell 96 secured to the combustor walls 60, 62, and a multiple of circumferentially distributed bulkhead heatshields segments 98 secured to the bulkhead support shell 96 around the central opening 92.

The annular hood 82 extends radially between, and is secured to, the forwardmost ends of the combustor walls 60, 62. The annular hood 82 includes a multiple of circumferentially distributed hood ports 94 that accommodate the respective fuel nozzle 86 and introduce air into the forward end of the combustion chamber 66 through a central opening 92. Each fuel nozzle 86 may be secured to the diffuser case module 64 and project through one of the hood ports 94 and through the central opening 92 within the respective fuel nozzle guide 90.

The forward assembly 80 introduces core combustion air into the forward end of the combustion chamber 66 while the remainder enters the outer annular plenum 76 and the inner annular plenum 78. The multiple of fuel nozzles 86 and surrounding structure generate a blended fuel-air mixture that supports combustion in the combustion chamber 66.

Opposite the forward assembly 80, the outer and inner support shells 68, 70 are mounted to a first row of Nozzle Guide Vanes (NGVs) 54A in the HPT 54. In one disclosed non-limiting embodiment, thirty-two (32) NGVs 54A are located immediately downstream of the combustor 56. The NGVs 54A in one disclosed non-limiting embodiment, are the first static vane structure upstream of a first turbine rotor in the turbine section 28. The NGVs 54A are static engine components which direct core airflow combustion gases onto the turbine blades of the first turbine rotor in the turbine section 28 to facilitate the conversion of pressure energy into kinetic energy. The core airflow combustion gases are also accelerated by the NGVs 54A because of their convergent shape and are typically given a “spin” or a “swirl” in the direction of turbine rotor rotation. The turbine rotor blades absorb this energy to drive the turbine rotor at high speed. The NGVs 54A generate bow waves along a leading edge 54L which may shorten combustor 56 service life.

With reference to FIG. 3, studs 100 extend from the liner panels 72, 74 to mount the liner panels 72, 74 to the respective shells 68, 70 with fasteners 102 such as nuts. That is, the studs 100 extend from the liner panels 72, 74 for passage through the respective shells 68, 70 to receive the fasteners 102. Impingement cooling holes 104 penetrate through the shells 68, 70 to allow a coolant from the respective annular plenums 76, 78 to enter cavities 106A, 106B (FIG. 4) formed in the combustor liner assemblies 60, 62 between the respective shells 68, 70 and liner panels 72, 74. Film cooling holes 108 penetrate each of the liner panels 72, 74 to allow the cooling air to pass from the cavities 106A, 106B along a cold side 110 of the liner panels 72, 74 to a hot side 112 of the liner 72, 74 to promote the formation of a film of cooling air over the hot side 112.

The cavities 106A, 106B are defined by back side features 114 such as rails that extend form the cold side 110 to contact the respective shells 68, 70 to form enclosed spaces in the liner assemblies 60, 62 when the liner panels 72, 74 are mounted to the shells 68, 70. It should be appreciated that back side features 114 other than rails such as posts, trip strips and others may alternatively or additionally extend from the cold side 110.

The cold side 110 of the liner panels 72, 74 includes a cold side coating 116 that in the disclosed non-limiting embodiment is a diffusion-aluminide coating that is not readily stripped from the liner panels 72, 74 without damage theretor. The hot side 112 of the liner panels 72, 74 includes a hot side coating 118 that in the disclosed non-limiting embodiment is a thermal barrier coating such as a nickel-aluminide.

With Reference to FIG. 4 in one disclosed non-limiting embodiment of a repair process, the thermal barrier coating is stripped in a localized area which is damaged to facilitate a variable polarity weld (step 200). Alternatively, the thermal barrier coating is stripped completely from the liner panels 72, 74. In one disclosed non-limiting embodiment, the thermal barrier coating is acid striped which removes the thermal barrier coating but does not effect the diffusion-aluminide coating from the cold side 110.

Alternatively still, the thermal barrier coating is not stripped as development trials have demonstrated that variable polarity welding facilitates local weld repair of damage on coated parts without the need to completely remove all coating. Variable polarity welding essentially locally strips the coating from the repair area during welding which prevents microstructural damage via coating/base metal interaction and does not damage any of the surrounding coating.

In step 210, the damaged area is then welded via Variable Polarity welding. Variable Polarity welding utilizes an unbalanced straight and reverse polarity weld current verses balanced A/C weld current. This eliminates the requirement of continuous high frequency current for arc stability and facilitates a low heat input weld process that improves weldability in thin sections sectioned and difficult to weld high temperature resistant hardenable Nickle-based super alloys.

In step 220, the damaged area or the entire hot side 112 is then recoated with a thermal barrier coating. The thermal barrier coating may be locally applied.

It should be appreciated that although a liner panel is in illustrated in the disclosed non-limiting embodiment, other components will also benefit from the disclosed repair.

It should be understood that relative positional terms such as “forward,” “aft,” “upper,” “lower,” “above,” “below,” and the like are with reference to the normal operational attitude of the vehicle and should not be considered otherwise limiting.

It should be understood that like reference numerals identify corresponding or similar elements throughout the several drawings. It should also be understood that although a particular component arrangement is disclosed in the illustrated embodiment, other arrangements will benefit herefrom.

Although particular step sequences are shown, described, and claimed, it should be understood that steps may be performed in any order, separated or combined unless otherwise indicated and will still benefit from the present disclosure.

The foregoing description is exemplary rather than defined by the limitations within. Various non-limiting embodiments are disclosed herein, however, one of ordinary skill in the art would recognize that various modifications and variations in light of the above teachings will fall within the scope of the appended claims. It is therefore to be understood that within the scope of the appended claims, the disclosure may be practiced other than as specifically described. For that reason the appended claims should be studied to determine true scope and content. 

What is claimed is:
 1. A liner panel for use in a combustor of a gas turbine engine comprising: a cold side having a backside feature; and a hot side with a variable polarity weld.
 2. The liner panel as recited in claim 1, wherein said hot side includes a thermal barrier coating.
 3. The liner panel as recited in claim 2, wherein said thermal barrier coating is a nickel-aluminide.
 4. The liner panel as recited in claim 2, wherein said thermal barrier coating is an aluminum oxide.
 5. The liner panel as recited in claim 1, wherein said cold side includes a coating.
 6. The liner panel as recited in claim 5, wherein said coating is a diffusion-aluminide.
 7. A method of repairing a component of a gas turbine engine comprising: welding a damaged area via variable polarity welding.
 8. The method as recited in claim 7, further comprising: stripping a thermal barrier coating prior to the welding.
 9. The method as recited in claim 8, further comprising: applying a thermal barrier coating after the welding.
 10. The method as recited in claim 7, further comprising: acid stripping a thermal barrier coating prior to the welding.
 11. The method as recited in claim 7, further comprising: stripping a thermal barrier coating in the damaged area prior to the welding.
 12. The method as recited in claim 11, further comprising: applying a thermal barrier coating in the damaged area after the welding.
 13. A liner panel for use in a combustor of a gas turbine engine comprising: a cold side having a coating; and a hot side with a variable polarity weld.
 14. The liner panel as recited in claim 13, wherein said hot side includes a thermal barrier coating.
 15. The liner panel as recited in claim 14, wherein said thermal barrier coating is a nickel-aluminide.
 16. The liner panel as recited in claim 13, wherein said cold side includes a coating.
 17. The liner panel as recited in claim 16, wherein said coating is a diffusion-aluminide. 